The present invention relates to abradable thermal barrier coatings, and more particularly relates to the use of such coatings for combustion turbine components such as turbine ring segments.
Metal components of combustion turbines are operated at very high temperatures and often require the use of thermal barrier coatings (TBCs). Conventional TBCs typically comprise a thin layer of zirconia. In many applications, the coatings must be erosion resistant and must also be abradable. For example, turbine ring seal segments which fit with tight tolerances against the tips of turbine blades must withstand erosion and must preferentially wear or abrade in order to reduce damage to the turbine blades.
In order to provide sufficient adherence to the underlying metal substrate, conventional TBCs are provided as relatively thin layers, e.g., less than 0.5 mm. This thickness is limited by the thermal expansion mismatch between the coating and metallic substrate. However, such thin layers limit the heat transfer characteristics of the coatings, and do not provide optimal erosion resistance and abrasion properties.
The goal of achieving improved gas turbine efficiency relies upon breakthroughs in several key technologies as well as enhancements to a broad range of current technologies. One of such key issues is to tightly control rotating blade tip clearance. This requires that turbine ring segments, also known as turbine heat shields or turbine outer seals, are able to absorb mechanical rubbing against rotating blade tips.
For closed loop steam cooled turbine ring segments, a thick thermal barrier coating of about 0.1 inch on the ring segment surface is required for rubbing purposes. The latest advanced gas turbine has a hot spot gas temperature of 2,800xc2x0 F. at the first stage ring segment. Under such a high thermal load, a TBC surface temperature of 2,400xc2x0 F. is expected. Thus, the conventional abradable TBC is no longer applicable because TBC has a limitation of maximum surface temperature up to 2,100xc2x0 F.
Electron beam physical vapor deposited thermal barrier coatings (EB-PVD TBCs) are a possible alternative solution for such high surface temperatures. However, EB-PVD TBCs are not very abradable and are not considered satisfactory for conventional turbine ring segment applications.
Friable graded insulation (FGI) comprising a filled honeycomb structure has been proposed as a possible solution to turbine ring segment abrasion. FGI materials are disclosed in U.S. patent application Ser. No. 09/261,721, which is incorporated herein by reference. The use of FGI as an effective abradable is based on the control of macroscopic porosity in the coating to deliver acceptable abradability. The coating consists of hollow ceramic spheres in a matrix of aluminum phosphate. The ability to bond this ceramic coating to a metallic substrate is made possible by the use of high temperature honeycomb alloy which is brazed to a metallic substrate. The honeycomb serves as a mechanical anchor for the FGI filler, and provides increased surface area for chemical bonding. However, one key issue relating to the practical use of FGI honeycomb coatings applications such as turbine ring segments is that the edges and corners of the ring segments are exposed to hot gas convection. Wrapping the filled honeycomb around the edges and corners presents distinct difficulties for manufacturing.
The present invention has been developed in view of the foregoing, and to address other deficiencies of the prior art.
The present invention provides a high temperature, thermally insulating and/or abradable composite coating system that may be used in gas turbine components such as ring seal segments and the like. The coating system includes a first composite thermal barrier coating covering a portion of the component, and a second deposited thermal barrier coating covering edge portions of the component.
The preferred first composite thermal barrier coating includes a composite material which comprises a metal base layer or substrate, a metallic honeycomb structure, and a ceramic filler material. The ceramic filler material preferably comprises hollow ceramic spheres within a phosphate matrix to provide high temperature capability and excellent thermal insulation. The resulting system is compliant and accommodates differential thermal strains between the ceramic and the metallic substrate material. The honeycomb/ceramic composite may optionally be overlaid with a ceramic layer to protect and insulate the metallic honeycomb.
The second deposited thermal barrier coating covers edge portions of the component, and preferably comprises a combination of zirconia and yttria, e.g., ZrO2-8 wt % Y2O3. The deposited thermal barrier edge coating is preferably applied by electron beam physical vapor deposition (EB-PVD) techniques. The EB-PVD ceramic preferably has a columnar microstructure which may provide improved strain tolerance. Under mechanical load, or thermal cycling, the ceramic columns produced by EB-PVD can move, both away from each other and towards each other, as strain cycles are applied to a component.
In addition to improved thermal properties, the present coating system displays excellent abradable properties. The honeycomb structure of the first composite coating provides good adhesion between the ceramic material and the underlying metallic substrate/component. By infiltrating the ceramic into the cells of the honeycomb during processing, the honeycomb provides additional mechanical anchoring to enhance ceramic to metal adhesion. The composite enables the use of relatively thick insulating coatings, e.g., on the order of 2 mm or more, to provide very high temperature protection to metallic hot section gas turbine parts.
The coating system in addition to providing adequate abradability also possesses excellent erosion resistance. For example, the ceramic on a ring seal segment should wear preferentially to the metal of a blade in the case of ring seal segment/blade tip rubbing. This property provides the capability to restrict blade tip clearances and to improve engine efficiencies without incurring the damage to blade tips that conventional TBC coatings cause in similar situations.
The present invention provides a more durable, cost effective thermal barrier coating system for use with ring seal segments, transitions, combustors, vane platforms, and the like.
An aspect of the present invention is to provide a thermal barrier coating system comprising a metal substrate, a first composite thermal barrier coating over a portion of the substrate, and a second deposited thermal barrier coating over at least an edge portion of the substrate adjacent a periphery of the first composite thermal barrier layer.
Another aspect of the present invention is to provide a method of making a composite thermal barrier coating. The method includes the steps of covering a portion of a metal substrate with a first composite thermal barrier coating, and depositing a second thermal barrier coating over at least an edge portion of the substrate adjacent a periphery of the first composite thermal barrier layer.
These and other aspects of the present invention will be more apparent from the following description.